A mount system for tiltably supporting a pylon assembly of a rotorcraft. First, second, third and fourth pylon links are each coupled between the pylon assembly and the airframe of the rotorcraft. The first pylon link has a first axis, the second pylon link has a second axis, the third pylon link has a third axis and the fourth pylon link has a fourth axis. Each of the axes intersects at a focal point located proximate a coaxial rotor system having counter-rotating upper and lower rotor assemblies such that the focal point provides a virtual pivot point about which the pylon assembly tilts to generate a control moment about a center of gravity of the rotorcraft that counteracts lateral and fore/aft moments generated by the upper and lower rotor assemblies during rotorcraft maneuvers.

A rotor system including a rotor hub, a plurality of rotor blades supported by the rotor hub, and a fairing mounted to the rotor hub. The fairing includes an external surface exposed to an external airflow and an internal surface defining an interior portion. A component system is mounted to the rotor hub in the interior portion. The component system includes a first heat generating component and a second heat generating component spaced from the first heat generating component. A cooling duct is arranged between the first heat generating component and the second heat generating component.

A method for predicting vibrations in an aircraft comprising an active vibration reduction system includes estimating a first vibration amplitude or frequency resulting from adjustments by the active vibration reduction system and the respective sensitivities of the aircraft depending on the flying state using a statistical mathematical process, recording a second vibration amplitude or frequency by a sensor, generating a pseudo-vibration profile by combining the first and second vibration amplitudes or frequencies, comparing the pseudo-vibration profile with a predefined target vibration profile, and outputting a signal when a specific threshold value has been exceeded.

Provided is a vibration control system for a compound helicopter with a rotor and a fixed wing. The fixed wing includes a movable flap that is mounted on a rear edge of the fixed wing. The vibration control system periodically moves the movable flap so as to periodically change lift of the fixed wing such that vibration aerodynamically generated by the fixed wing is in anti-phase with vibration caused by rotation of the rotor.

A vibration attenuator for a rotor of an aircraft has a first ring with an eccentric weight, a coaxial second ring with an eccentric weight, and a central ring coaxial with the first and second rings and located therebetween. A first post extends from the first ring toward the central ring and is received in a first arcuate groove formed on the central ring, whereas a second post extends from the second ring toward the central ring and is received in a second groove formed on the central ring. A motor is configured for driving the central ring in rotation about the axis relative to the motor. The grooves are equal in length, and a center of the first groove is located on an opposite side from a center of the second groove. Rotation of the central ring by the motor causes rotation of the first and second rings.

A noise reduction system for air mobility, may include a plurality of propeller modules provided on the air mobility and each including an inverter, a motor electrically connected to the inverter, and a propeller mounted to the motor; and a control unit configured to perform PWM control of power to be provided to the inverters to reduce noise attributable to carrier frequencies generated by the plurality of inverters or reduce noise attributable to fundamental frequencies generated by the plurality of motors.

Systems and method for active vibration control on an aircraft. An active vibration control system (AVCS) is configured for an aircraft having an aircraft structure and a gun. The AVCS includes at least one control sensor on the aircraft, at least one force generator on the aircraft, and at least one controller in electronic communication with the sensor and the force generator. The controller is configured for determining, using the at least one control sensor, force generating commands for controlling vibrations acting on the aircraft structure, sending the force generating commands to the at least one force generator, causing the at least one force generator to produce a vibration canceling force, determining that the gun is firing, and in response to determining that the gun is firing, determining different force generating commands and sending the different force generating commands to the at least one force generator.

One example of a mount for a rotorcraft comprises a structural support member, a bracket, and an elastomer. The bracket is configured to attach to a component of the rotorcraft. The component of the rotorcraft produces vibrations at a first frequency. The structural support member configured to transfer a weight of the component of a rotorcraft to an airframe of the rotorcraft. A rotor system of the aircraft vibrates the airframe of the rotorcraft at a second frequency. The elastomer is located between a structural support member and a bracket. The elastomer is configured to attenuate noise caused by the vibrations at the first frequency by isolating the vibrations at the first frequency from reaching the airframe of the rotorcraft while the airframe vibrates at the second frequency.

A linear vibration damper and/or absorber includes a centre shaft (12) having bearing regions (18) and friction contact regions (20), and a housing (2) including finger assemblies (22) which are mounted with a small radial clearance for accurate location on the bearing regions (18) for axial displacement with respect to the centre shaft (12) along a central axis (X), the finger assemblies (22) each including resilient fingers (38) which extend axially from respective body sections (26) and have contact faces (40) which resiliently engage, i.e. are pressed by the resilience of the fingers (40) into contact with, friction surfaces (20) of the contact regions of the centre shaft (12), whereby relative linear displacement between the centre shaft (12) and the housing (2) is opposed by frictional contact between the friction surfaces (20) and the contact faces (40).

In an embodiment, an aircraft includes a fuselage and a first support boom extending from the fuselage and having a first boom thickness. The aircraft also includes a first propulsion assembly coupled to the first support boom. The first propulsion assembly includes a first rotor hub and first rotor blades extending from the first rotor hub and operable to rotate in a first rotor plane with the first rotor hub. The aircraft also includes a first separation between the first rotor plane and the first support boom of not less than approximately 1.5 times the first boom thickness.